1

ZERO BOIL-OFF SYSTEM TESTING

D.W. Plachta1 and W.L. Johnson1, J.R Feller2

1NASA Glenn Research Center 2NASA Ames Research Center

Abstract

Cryogenic propellants such as liquid hydrogen (LH2) and liquid oxygen (LO2) are a part

of NASA’s future space exploration plans due to their high specific impulse for rocket

motors of upper stages. However, the low storage temperatures of LH2 and LO2 cause

substantial boil-off losses for long duration missions. These losses can be eliminated by

incorporating high performance cryocooler technology to intercept heat load to the

propellant tanks and modulating the cryocooler temperature to control tank pressure. The

technology being developed by NASA is the reverse turbo-Brayton cycle cryocooler and

its integration to the propellant tank through a distributed cooling tubing network coupled

to the tank wall. This configuration was recently tested at NASA Glenn Research Center

in a vacuum chamber and cryoshroud that simulated the essential thermal aspects of low

Earth orbit, its vacuum and temperature. This test series established that the active

cooling system integrated with the propellant tank eliminated boil-off and robustly

controlled tank pressure.

Introduction

To expand human presence into the solar system and onto the surface of Mars, NASA is

considering high-specific-impulse propellant combinations, such as LH2 and LO2, for

orbiting depots, orbit transfer stages, and for Mars surface. However, for volumetric

considerations, cryogenic propellants are stored as liquids at extremely low temperatures.

The heat radiated to the spacecraft from both the Sun and any other celestial body the

spacecraft may be near (such as Earth, the moon, Mars, etc.) and heat conducted down to

the storage tanks from other sources on the spacecraft cause LH2 and LO2 to pressurize

and boil off (change state from liquid to gas). If this is left to its own devices, the storage

tanks will over-pressurize; thus they must vent some of the vaporized liquid, resulting in

less propellant remaining available for propulsion. Because mission architecture loiter

periods are projected to be months long [1], the vented vapor losses will be substantial.

To allow for these losses during the long space missions envisioned, the stage would

need to carry excess propellant which would be very heavy. Alternatively, NASA could

use thick walled propellant tanks that operate at high pressure, but the additional mass of

the heavier tanks would also be prohibitive. The application of Zero Boil-Off (ZBO)

technology to prevent vaporization and keep storage tanks reasonably sized and low

weight will enable missions to store adequate propellant quantities for long periods of

time. Development work and testing on this concept using distributive cooling has been

2

on going at NASA since 2007, with the Cryogenic Boil-Off Reduction System

activities[2],[3],[4]. Analysis of this ZBO concept applied to liquid oxygen tanks

predicts that it will reduce mass for missions in low Earth orbit (LEO) that have loiter

periods greater than 1 week.[5] The preferred distributed cooling system utilizes the

reverse turbo-Brayton cycle cryocooler and the circulator that is inherent to it. This

concept and associate technology was demonstrated in a series of 10 tests performed at

NASA Glenn Research Center’s Small Multi-Purpose Research Facility (SMiRF). Three

“passive,” with the cryocooler system off, and seven that are “active,” with the

cryocooler system operational. The test series included tests performed at roughly 90%

full and 25% full, and demonstrations of cryocooler excess cooling capacity countered

with tests at reduced cryocooler capacity. The test series established that the prescribed

cryocooler integration system eliminated boil-off and robustly controlled tank pressure.

Objectives

The purpose of the test was to demonstrate the performance of a flight

representative liquid oxygen (LO2) ZBO system. Given the lack of micro-gravity

concerns with the active cooling system or on the unvented propellant, this

demonstration, prepares the ZBO concept for flight with minor additional development

required beyond scaling of components. To achieve this demonstration, three main

objectives were identified.

The first objective was to demonstrate robust zero boil-off storage of liquid

oxygen. This requires a demonstration of the ability of the active cooling system to

control and modulate tank pressure over an extended period of time. Besides achieving

ZBO, a demonstration of the active cooling system dropping tank pressure is significant

as it indicates the system has performance margin to account for various uncertainties in

the design. Inherent in robust ZBO is the ability to model tank pressurization and

depressurization. Because the distributed cooling system in the ZBO design promises to

reduce thermal stratification, a simple homogenous model might be accurate. This is

planned along with a comparison test, with the cryocooler off, to create a mapping of

tank pressure versus net heat removed and added. The second objective was to determine

the cryocooler’s ability to eliminate boil-off at a low fill level. This is needed for

propellant depot or upper stage mission concepts that have multiple transfers or

propellant burns that need to be operational at fairly low tank fill levels. Low fill levels

increase thermal gradients in the tank and the goal of this second objective is see if the

ZBO system can minimize those increased thermal gradients and still maintain tank

pressure control and ZBO. The third test objective was to validate the scaling model

developed [5, p. 66] that predicts ZBO thermal elements, such as multi-layer insulation

(MLI), the cryocooler, radiator, and solar arrays, reduce mass compared to always boiling

passive only propellant storage, with MLI only, for loiter periods in LEO of just over one

week. Given that this prediction is based on analysis, verification is required to ensure

mission architecture consideration of the active cooling concept is done appropriately.

Although the mission architectures are interested in LO2, the high testing costs

associated with it caused us to use liquid nitrogen (LN2) as a LO2 simulant. A nominal

pump fed propulsion system LO2 storage pressure was assumed to be 173kPa (25 psi),

which corresponds to a saturated LO2 temperature of 95.6 K. To accomplish this, the test

3

tank was filled with LN2 and pressurized to 565 kPa (82 psia). This is the saturation

pressure when LN2’s saturation temperature (95.6 K) is equal to that of LO2.

Hardware Description

Facility Overview

The experiment was conducted at NASA GRC’s Small Multi-Purpose Research Facility

(SMiRF) [6]. SMiRF provided the two main aspects of a LEO simulation test—the

vacuum of space and the temperatures of Low Earth Orbit (LEO). The SMiRF facility

utilizes a cylindrical vacuum chamber with elliptical heads and achieved an average

vacuum of 1x10-6 torr vacuum throughout the test series. To accurately simulate the

space environment, an optically dense flat black painted thermal shroud, or cryoshroud, is

fitted closely within the vacuum chamber walls. The cryoshroud is operated as a constant

density closed loop GN2 heating/cooling system which was operated at 220 K +/- 3 K for

9 of the 10 tests. The shroud reduces the maximum allowable size of the test article to a

diameter 1.5 m and overall length of 2 m. The test article is shown in Figure 1 attached

to the vacuum chamber lid at SMiRF and being lowered into the chamber.

Liquid Nitrogen Test Tank

The test tank is stainless steel with a diameter of 1.2 m and a 4.7 mm wall thickness that

is 1.2 m3 in volume. The tank height is 1.4 m and the length to diameter ratio is 1.15. The

4

domes are 2:1 elliptical head domes with a 0.7 m long cylindrical section in between. The

tank was attached to the six struts and via three attachment plates. The struts were 0.38 m

long, having a tapered geometry with a maximum outer diameter of 19 mm and a 0.82

mm wall thickness. These titanium struts have spherical rod end bearings at both ends.

The tank has twelve heaters attached to its outer diameter at the bottom part of the tank

cylinder, which allowed for rapid warm-up of the tank between tests.

The propellant tank maximum operating pressure was 620 kPa (90 psia). The nominal

operating pressure was 565 kPa (82 psia).

At the top port of the tank, used for tank venting, a cooling strap was coupled as close to

the tank as possible, to reduce the vent line temperature. This was designed and installed

because of a pre-test finite element thermal model analysis that indicated a hot spot at the

tank top. A model of the tank with the tank penetrations used and all the associated heat

leak paths used in the thermal analysis is shown in Figure 2.

5

Support Ring

The support ring is suspended from the SMiRF chamber lid by three cables. The ring

supports the tank, the cryocooler, and the radiator and is made from stainless steel. The

layout of the components configured in the ring is shown in Figure 3.

6

Figure 3. Cryocooler Layout in Support Ring (top view)

Radiator

The radiator was designed to remove 400 W of heat at 300 K. It is a curved aluminum

panel that is 3 mm thick. Attached to this panel were four horizontal ammonia heat pipes

of 9 mm diameter. At the end of the radiator panel is the evaporator plate where the

cryocooler hot interface was attached. The radiator was insulated with 10 layers of MLI

on its inside surface, to ensure that the vast majority of the heat radiates from its outer

diameter surface. Its outer surface was painted white with Aeroglaze A276 paint with a

measured emissivity was 0.935.

Insulation

In order to minimize the heat load on the acreage of the tank, a multilayer insulation

system was required. Two MLI blankets were constructed, each with 38 layers (including

outer covers) for 75 total reflector layers. The Mylar used was 0.63 mm (0.25 mil)

double aluminized Mylar (DAM). Each layer was separated by 2 sheets of Dacron B2A

polyester netting. The blankets were vented through the seams with a 1% open area in

the outer 2-mil cloth reinforced layer of Mylar. The seams were butted together and

adjoined with Mylar (2.5 cm wide) tape every fifth layer. The as built blanket layer

density was 24 layers/cm.

7

A 12 mm strip of Cryo-Lite [7] was wrapped around the base of each penetration, and the

tank and penetration MLI butted against it, as recommended in reference [8]. The tank

penetrations were wrapped with 15 layers of MLI, again with two Dacron netting spacers

used between each Mylar layer.

Broad Area Cooling

The broad area cooling tubing network consisted of 10 tubes, 5 supply and 5 return that

went vertically down the tank wall (see Figure 2). Each tube was 0.64 cm (¼ inch)

tubing with a 0.89 mm (0.035 inch) wall thickness distributed evenly around the tank.

The tubes were spaced 36 degrees apart from each other and were coupled together at the

tank top by using two manifolds, 1.3 cm diameter.

The cooling tubes were epoxied on one side down the length of each tube. In addition,

the tubes were spot welded to the tank at every 0.30 m of tube length. This configuration

resulted in a tube-to-tank thermal effectiveness of 0.9, which was satisfactory. This

structural and thermal concept was settled upon after numerous epoxy configurations

were tested with LN2, to evaluate the tube epoxy. The epoxy selected was 3M’s Scotch-

Weld 2216. By itself, however, the epoxy did not have enough strength to secure the tube

tight to the tank wall, due to contraction of the tank wall at liquid nitrogen temperatures.

Cryocooler

8

The schematic of the reverse turbo-Brayton cryocooler cycle used in the test is shown in

Figure 4, with instrumentation. The cycle gas is neon. A single compressor compresses

the neon, causing it to flow continuously through the system. The heat of compression

and any losses within the compressor are rejected at the radiator mounting plate, which

also rejects any losses associated with the electronics. The high-pressure gas flows

through both recuperators and then is expanded through the turbo alternator and flows to

our broad area cooling system. The cryocooler/circulator is a modification of an existing

2-stage cryocooler [9] that was designed and built by Creare that was originally based on

the NICMOS cryocooler used on Hubble Space Telescope. It was modified by (1)

eliminating the second-stage turbo-alternator and recuperator; (2) replacing the

commercial compressor inlet filter and aftercooler with flight-like versions; (3) altering

the compressor flow passages for lower flow rates; and (4) repackaging the cryocooler

assembly and reconfiguring the tubing, valves, and fittings to interface with the NASA-

provided distributed cooling network. A three-dimensional model of the cryocooler is

shown in Figure 5.

9

Figure 5. A 3D model of the reverse turbo-Brayton cryocooler assembled into the aluminum

channel structure.

The cryocooler’s working fluid is neon at ~2 atm, with a nominal flow rate of 2 g/s.

The cooling capacity specified is 15 W at a load temperature of 77 K, however the lift

was over 20 W at the nominal cryocooler return temperature of 98.4 K, used for the ZBO

tests. The cryocooler was operated by setting the return temperature to a user specified

value. There was no direct feedback to tank pressure, so the cryocooler temperature set

point was finely adjusted until tank pressure was steady. For the pressurization tests, the

compressor input power was varied which changed the cryocooler lift from ~3 W to over

20 W at neon mass flow rates between ~1.5 g/s to 2.5 g/s.

Heat was generated at the compressor and aftercooler, which are both mounted on a

common mounting plate. This plate was thermally coupled to the radiator, where the heat

was rejected to the cryoshroud. The design heat rejection temperature of the cryocooler is

between 270 and 300 K.

Instrumentation

The recorded cryocooler data is shown in Figure 6 and includes redundant sensors at the

inlet and exit conditions at each main component in the system. The BAC temperatures,

T6 and T7, as shown in the schematic, were calibrated platinum resistant thermometry

probes (+/- 0.04 K) that measured the neon temperature in the tube. The cryocooler

10

system mass flow rate is found from the fluid supply and return conditions of the turbo-

alternator, its rotational speed, and its physical characteristics.

In addition to the cryocooler system, the propellant tank and the rest of the ZBO test

article was highly instrumented. Measurements used for conducting the test series

included tank pressure, vacuum chamber pressure, tank liquid and wall temperatures,

insulation temperatures, cryoshroud temperatures, BAC temperatures, and tank boil-off

flow rate. The types of instruments, numbers, and their locations are shown in Table 1.

Table 1. ZBO Instrumentation.

Location Count SD/TC Notes – Purpose

LN2 Temperature

Sensor Rake

8 8/0 Liquid temperature and liquid level indication.

Key sensors at 96.9, 87.2, and 28.4 % full.

Tank Wall 13 12/1 Exterior tank temperatures at top, bottom, and

between cooling loops.

BAC System 28 21/7 Measure BAC system temperatures (cooling

tubes, manifolds, and thermal strap)

Penetrations 16 6/10 Two at warm and two at cold end of vent,

fill/drain, and cap probe. Used for heat leak

calc’s

Struts 26 2/24 Two at warm and two at cold ends. Heat leak

calculations.

Radiator 25 0/25 Map radiator performance.

MLI 11 0/11 Determine MLI temperature profile.

Supports/cabling 12 0/12 Used to find misc. heat leak through wire

bundles & suspension hardware.

Cryoshroud 18 0/18 Boundary temperature definition and control.

Tank Pressure 2 NA Measure and control tank pressure. Range of

sensors were 0-50 and 0-100 psia.

Boil-off Flow 4 NA Mass flowmeters used to measure boil-off rates

Tank/Strut

Heaters

14 NA Warm up tank, warm liquid, and set warm

boundary temperature on struts

Key—SD refers to a silicon diode; TC is a thermocouple.

11

Test Plan

The test was planned to meet the major test requirements of establishing the baseline heat

leak, achieving zero boil-off by controlling tank pressure for an extended period of time,

and robustly controlling pressure by dropping tank pressure using the cryocooler. To

accomplish this, the test tank was filled with LN2 and then pressurized to 565 kPa. This

is the saturation pressure when LN2’s saturation temperature (95.6 K) is equal to that of

LO2 at 172 kPa, the nominal pressure fed propulsion system tank supply pressure. Tank

sidewall heat was applied to facilitate the pressure rise. The fluid liquid level was set to

~95% full. The environmental conditions were set using the cryoshroud, which was set

to 220 K, a representative low Earth orbit temperature, and the vacuum pumps, which

evacuated the chamber to 1x10-6 torr. These conditions remained constant for the vast

majority of the test series.

Two types of tests were conducted, steady state tests and pressurization tests. Steady

state tests were as the name applies, performed until the steady state criteria was

established, with the most important one being that the MLI temperatures did not vary

more than 0.55 K in a six hour period. Pressurization (including depressurization) tests

were performed to understand the tank heating rate effect on tank pressure.

Pressurization tests were conducted overnight, to ensure enough time for sufficient

pressure variation.

Table 2 shows the tests conducted.

Fill Level Type Purpose

Test 1 95% Passive boil-off Find tank heat leak

Test 2 95% Passive Pressurization Find tank pressure rise rate

Test 3 95% ZBO Achieve ZBO; collect data

Test 4 95% ZBO high power Find robustness of ZBO system

Test 5 95% ZBO low power More data to map pressurization rate with

cooler power

Test 6 95% ZBO destratification Find tank pressure rise rate with tank heat

added while at ZBO

Test 7 95% ZBO high power 2 More data to map pressurization rate with

cooler power

Test 8 25% ZBO Achieve ZBO; collect data

Test 9 25% ZBO high power More data to map pressurization rate with

cooler power

Test 10 95% Passive boil-off with

cryoshroud set to

300K

Additional MLI data point for tank applied

system

The cryocooler was operated continuously from Test 3 through Test 9 for 19 days and

during that time, the test tank was not vented.

Results

Summary of Component Performance

12

There were several components that were analyzed pre-test for their performance, as

shown in Table 3, and are compared here with their actual performance. Also, comments

on further needed developments for flight are included.

Table 3 Component Performance Pre-Test Analysis Result

BAC Network Temp Increase < 5 K Temp gradient 3.8 K (max)

BAC Pressure Drop <4.1 kPa Not Accurately measured

BAC tube-tank Gradient 0.1 K 0.5 K

BAC Effectiveness 99% 90%

Cryocooler % of Carnot 10.6% 10.6%

Parasitic loss No model 4.2 W (ave)

Broad Area Cooling

The Broad Area Cooling system design performed well. The two main indicators of that

were the temperature gradient and pressure drop. The temperature gradient from the

exterior tank top to bottom was 3.8 K in Test 3, a large reduction from Test 1, where the

same gradient was 10.2 K. Also, the pressure drop in the system, ∆P2, across the BAC

network, was 1.7 kPa (0.25 psi), less than the 4.1 kPa expected. This is for a tubing

network that was 4.2 m long on the tank, plus the manifolds and 1 m long supply and

return hoses. The low thermal gradient from tank top to bottom contrasts dramatically

with the 2003 Advanced ZBO Demonstration Test[10], with its “S” link (20 cm long) and

heat pipe (80 cm long and submerged in the LN2 tank) coupling of the pulse tube

cryocooler to the LN2 tank. That thermal gradient was 6.9 K, which would have

increased significantly if the cryocooler was not placed adjacent to the tank.

The tube-to-tank thermal gradient was 0.5K, higher than the pre-test model indicated.

However, the tube-to-tank heat exchanger effectiveness analysis performed, which used

the tank fluid, wall, and BAC tube temperatures in its calculations, found the tube-on-

tank thermal effectiveness was 90%. The effect of this effectiveness on cryocooler input

power was found to be minimal, when compared to a 100% effectiveness. The impact

for Test 5, for example, was an increase in cryocooler input power of 0.5 W, a 0.4%

increase. A configuration with a 70% effective BAC would require 1.4% increase in

input power.

Besides the outstanding thermal aspect to the BAC, the fluid flow and structural results

with a cryogen were also promising. The low thermal gradients also indicate that the

fluid flow was evenly distributed through the distributed cooling network that had no trim

valves or orifices to even out flow. This indicates that the manifolds effectively

distributed the neon gas to each of the five cooling tubes. In addition, the thermal data

and the post-test destructive investigation shows that the spot weld and epoxy attachment

method worked properly in a system configuration, under vacuum and with cryogenic

nitrogen in the tank.

Despite the performance of the BAC tube-on-tank network, further work is needed for a

flight application. The design used was developed by bench testing different attachment

methods to a large, thick-walled steel pipe, which was then dipped into liquid nitrogen, to

13

determine if the tube would separate from the pipe. The next step in flight preparation is

a thermal and structural optimization with a flight-weight tank.

Cryocooler

The cryocooler performed nominally. For Test 3, which had relatively low input power,

the cryocooler’s efficiency was 10.6% of Carnot. For Test 4, the high power test, the

cryocooler’s efficiency was 12% of Carnot. Both of these cryocooler efficiencies fall on

the curve created in the Creare bench test and are nominal performances for flight

cryocoolers operating at these temperatures and heat loads.

The next step for flight preparedness for ZBO cryocoolers is improvements to the

cryocooler control system, to develop an algorithm to control tank pressure. It also needs

an improved method to operate at a set power level. Such improvements will give the

user more flexibility to control tank pressure robustly—to drop pressure in peak power

periods and allow it to increase in periods with solar eclipse. If this is done, testing

shows that the power can be effectively stored by modulating the tank pressure.

Parasitic Loss

The average parasitic loss for the test series was quite high at 4.2 W. Post-test studies

point to the high temperature and heat leak of the cryocooler return manifold as the

primary cause. The average measured emissivity for this Mylar tape was 0.67, much

higher than expected. If the tape had been low emissivity tape, with a polished aluminum

finish that has an emissivity of 0.03, like that of the double aluminized Mylar used in the

tank insulation, along with other improvements, analysis shows the parasitic loss would

have dropped to approximately 1.2 W to 1.5 W. This improves the ZBO system

efficiency from 5-9 to 7.5 to 11% of Carnot.

Parasitic heat leak estimation for a flight cannot be realistically estimated without a flight

configuration and associated thermal model. It is clear from the results of this test that a

thermal design minimizing the parasitic loss is mandatory. This requires thorough

understanding of MLI performance on the small tubes used in the broad area cooling

network. There are few references with data on MLI with small tubes.

Revisiting Test Objectives

This section will discuss how well the main test objectives were met.

Objective 1: Robust Tank Pressure Control

Robust tank pressure control was successfully demonstrated in this test series. This was

the first ZBO test with a distributed cooling system in which the cryocooler temperature

was used to modulate tank pressure. ZBO, cryo storage without venting, was

14

demonstrated over an extended 19 day period. The cryocooler was used to drop tank

pressure at two different power settings (see Figure 6). The tank pressure dropped at a

rate consistent with that predicted by a uniform temperature pressurization model.

Effectively, the propellant tank was used like a battery, to store power. This could be

very useful for in-space operation which will experience regular solar eclipses.

Additionally, the pressurization of a passive tank with no powered cryocooler (Test 2)

was compared with the pressurization of an actively cooled ZBO tank with heat added

(Test 6). For approximately the same duration, Test 2 tank pressure increased 36.2 kPa

while Test 6 tank pressure increased just 1.3 kPa. Adjusted per Watt of tank heat leak,

the rate of pressurization for Test 2 is 0.58 while Test 6 is 0.067 kPa/hr/W. That is, the

pressure rise rate of an effectively de-stratified cryogen is just 12 % of the stratified fluid.

Objective 2: ZBO at Low Fill Level

The cryocooler system was also used to achieve ZBO after a tank drain to approximately

25% full. This fill level, which causes increased thermal stratification due to the larger

ullage space, will occur in-space for any cryogenic propellant stage that undergoes

multiple engine burns or for a propellant depot after multiple propellant transfers to

cryogenic upper stages. This increased stratification cause tank pressurization rates to

increase, leading to more frequent vent cycles and increased boil-off losses. However,

the cryocooler broad area cooling system kept the tank top temperatures from increasing

significantly, from 98.7 at high fill level to just 98.9 K during the low fill level test.

These temperatures were much lower than Test 1, when the tank top temperature was

105.2 K.

The input power required to maintain steady state ZBO at this fill level was 145.9 W, 0.9

W higher than that in Test 3, a very slight increase. The BAC design proved to be more

than adequate at reducing the tank top temperatures to keep the tank at ZBO. Also, it

15

appears that the vent line cooling strap, included to keep tank lid temperatures low, was

not needed. The heat removed by it was less than its design and the temperatures it saw

were less than expected. Thus, the low fill level storage test requirement did not pose a

significant challenge to the active cooling system.

Objective 3: Validation of Scaling Study

In the Scaling Study [5], an Excel based Cryogenic Analysis Tool (CAT) was created to

predict cryogenic thermal control system performance and to estimate the in-space loiter

time at which active thermal control mass is equal to the passive thermal control mass,

including tank boil-off losses for liquid oxygen propellant tank storage. Following the

LOX ZBO tests and an analysis of the results CAT was updated using the design data and

the thermal system mass comparison was revisited. It was found that for a full scale 7.5

m diameter (182.6 m3) liquid oxygen tank with a nominal heat leak of 318 watts (through

75 layers of traditional MLI, structure, and penetrations), that the cryocooler system dry

mass increased 6.5% from our initial estimates, with the difference in total thermal

control system mass being 95 kg (out of 1456 kg). The updated model shifted the mass

breakeven in space loiter duration point from 7.2 to 8.0 days (see figure 7). Mission

durations longer than this should have this ZBO concept in their trade studies for

consideration. This shows that, although additional heat is added to the system due to

parasitic and integration losses that were not accounted for in previous versions, the

turbo-Brayton cryocooler system mass does not significantly increase, lending support to

the scalability of the LOX ZBO test concept.

16

LOX ZBO Test Summary

The LOX ZBO test was the first of its kind that demonstrated robust tank pressure control

using the cryocooler system to maintain and drop tank pressure without venting. The

tank pressure was maintained without venting for a 19 day period. The tank

stratification, which causes highly increased tank pressurization rates for unvented and

unmixed cryogenic tanks, was cut by 88% using the broad area cooling system, adjusted

for tank heating rate. Robust tank pressure control was also demonstrated at a low fill

level, which causes additional tank stratification. In this test, the broad area cooling

system minimized tank and fluid temperatures increases and ZBO was achieved with

virtually the same cryocooler input power as that at a high fill level. The tube-on-tank

broad area cooling system effectively prevented thermal stratifications within the tank

while being external to the tank and without introducing parasitic heat loads to the tank.

Because of these results, it is clear that an internal tank mixer, with its associated heat and

inherent risk to configurations with cryogenic propellants, is not required when the active

cooling system is operational.

The full ability of the cryocooler system was demonstrated. Tank pressure was

controlled to within +/- 0.1 psi using the active cooling system. Also, tank pressure was

decreased at controlled rates with the cryocooler operating at different levels of excess

capacity. This variation of the cryocooler’s input power and its ability to drop cryogenic

propellant tank pressure could reduce or eliminate the cryocooler stored input power

requirement, assuming solar arrays provide the power. Batteries are typically required

due to solar eclipses occurring during orbits in low Earth orbit.

The thermal results of the test series have been used to validate the scaling study analysis,

which has predicted large mass savings for applying ZBO for cryogenic upper stages or

depots exposed to long loiter periods in low Earth orbit. While the test found that

parasitic losses increase the cryocooler system mass and the passive to ZBO break-even

point was slightly longer, the many assumptions of the broad area cooling system and the

reverse turbo-Brayton cycle operation used in the modeling effort were confirmed.

For a potential flight application, this test series has advanced the technology

significantly, reducing the risk for future flight projects. The documented integrated

performance of four main components coupled to a cryogenic tank has increased the

confidence of this concept for flight. Work remains, particularly on optimizing the tube-

on-tank design and cryocooler integration parasitic design and analysis.

Acknowledgements

This work was funded and managed by NASA’s Space Technology Mission Directorate

under the Cryostat program.

17

Bibliography:

[1] Mars Architecture Steering Committee, "Human Exploration of Mars Design

Reference Architecture 5.0," NASA/SP–2009–566, page 3, 2009.

[2] D. W. Plachta, R. Christie and J. Feller, "Cryogenic Boil-Off Reduction System,"

AIP, New York, 2008.

[3] J. R. Feller, L. Salerno, A. Kashani and B. Helvenstein, "Distributed Cooling

Techniques for Cryogenic Boil-Off Reduction Systems," Kluwer

Academic/Plenum Publishers, New York, 2009.

[4] J. R. Feller, D. W. Plachta, G. Mills and C. McClean, "Demonstration of a

Cryogenic Boil-Off Reduction System Employing an Actively Cooled Thermal

Radiation Shield," Kluwer Academic/Plenum Publishers, New York, 2010.

[5] D. Plachta, "Cryogenic Boil-Off Reduction System Scaling Study," Elsevier,

www.elsevier.com/locate/cryogenics, 2014, Vol. 60, pages 62-68.

[6] J. Jurns and M. Kudlac, "NASA Glenn Research Center Creek Road Complex,"

Cryogenics Vol. 46, Pages 98-104, Elsevier, New York, 2006.

[7] "Lydall Performance Materials, Cryogenic Products, Cryo-Lite Technical Data

Sheet," 2011. [Online]. Available:

http://www.lydallthermal.com/what/CryogenicProducts.shtml. [Accessed July

2015].

[8] W. Johnson and A. Kelly, "Two Dimensional Heat Transfer around Penetrations in

Multilayer Insulation," NASA TP 2012-216315, page 50, 2012.

[9] M. Zagarola, "Demonstration of a Two-Stage Turbo-Brayton Cryocooler for Space

Application," Kluwer Academic/Plenum , Cryocoolers 15, New York, 2009.

[10] D. Plachta, "Results of an Advanced Development Zero Boil-Off Cryogenic

Propellant Test," NASA TM 2004-213390, 2004.